[0001] The present disclosure relates to cooling of engine accessories of a gas turbine
engine.
[0002] Conventionally the accessory gearbox of an aircraft gas turbine engine is mounted
within the outer engine nacelle in a location beneath the engine. The gearbox is connected
to the engine core by a radial drive shaft. The gearbox provides power to other accessories
such as an auxiliary generator and pumps for hydraulic fluid, fuel, oil etc.
[0003] Although effective in that the gearbox is kept away from the hot environment of the
engine core, the location of the gearbox within the outer nacelle is disadvantageous
in that it requires a relatively significant amount of space within the nacelle, which
can increase the overall diameter of the nacelle, leading to weight and drag increase
and adverse specific fuel consumption.
[0004] In alternative arrangements, it is possible to locate the accessory gearbox and other
accessories in an engine zone directly outside the casings which surround the engine
core. The accessories are mounted to the casings by brackets, short spars or bosses.
However, the high temperature environment near the engine core can produce thermal
loads which can reduce component reliability. For example, in some engines the zone
can operate at air temperatures above 300°C, which is above the thermal capability
of many accessory gearbox materials. Therefore to maintain acceptable surrounding
air temperatures for the gearbox, a ventilation system may be installed.
[0005] The present disclosure is at least partly based on a realisation that a possible
source of cooling air in turbofan engines is the bypass airflow through the bypass
duct of the engine. This airflow is typically at temperatures below 100°. One option,
therefore, is to form simple ventilation holes in the aerodynamic cowl which surrounds
the engine zone so that a small portion of the bypass airflow enters the zone to surround
the accessory gearbox. However, in high bypass ratio engines, particularly at low
engine power conditions, a low fan pressure ratio reduces the pressure difference
across the ventilation holes and thereby limits the amount of the bypass flow that
can be diverted in this way to cool the gearbox. This in turn limits the capability
of the ventilation system.
[0006] Thus according to a first aspect there is provided a gas turbine engine for an aircraft,
the engine including:
an engine core comprising a turbine, a compressor, and a core shaft connecting the
turbine to the compressor;
core casings surrounding the engine core;
an aerodynamic cowl which surrounds the core casings;
a propulsive fan located upstream of the engine core, the fan generating a core airflow
which enters the core engine and a bypass airflow which enters a bypass duct surrounding
the aerodynamic cowl; and
one or more engine accessories mounted in a space between the core casings and the
aerodynamic cowl;
wherein the gas turbine engine further includes:
one or more ventilation inlets which receive a portion of the bypass airflow as a
cooling flow for the engine accessories, the ventilation inlets being configured to
convert at least a portion of the kinetic energy of the bypass flow received therein
into a pressure rise such that the pressure of the cooling flow is increased relative
to the pressure of the bypass airflow; and
a manifold in fluid communication with the ventilation inlets to collect the cooling
flow therefrom, the manifold having plural exhaust holes therefrom through which the
cooling flow leaves the manifold to impinge upon and thereby cool the engine accessories.
[0007] Advantageously, by increasing the pressure of the cooling flow in this way, and directing
it to where it is needed using the exhaust holes of the manifold, the problem of a
low fan pressure ratio can be overcome such that even at low engine power conditions
sufficient cooling of the engine accessories can be achieved.
[0008] Optional features of the present disclosure will now be set out. These are applicable
singly or in any combination with any aspect of the present disclosure.
[0009] The one or more ventilation inlets may convert at least 50% of the kinetic energy
of the portion of the bypass airflow received therein into a pressure rise. Preferably
they may substantially stagnate the portion of the bypass airflow received therein.
[0010] The gas turbine engine may further include a control unit which is operable to regulate
the amount of cooling flow collected by the manifold from the ventilation inlets.
In this way, the amount of cooling flow can be adapted to different flight conditions,
reducing losses and improving engine efficiency. For example, the control unit can
be under the control of an engine electronic controller (EEC) which manages the operation
of the engine as a whole. In contrast, simple ventilation holes formed in the aerodynamic
cowl would not allow the amount of cooling flow to be adapted in this way so that,
for example, holes sized for the maximum power take-off condition where the engine
core operating temperatures are at their highest would tend to provide larger than
necessary ventilation flows at engine cruise, negatively impacting the performance
of the engine and increasing the engine specific fuel consumption.
[0011] The exhaust holes may be arranged such that the cooling flow impinging upon the engine
accessories forms a continuous film of cooling air around the accessories.
[0012] Conveniently, the exhaust holes may be formed into two lines which extend substantially
parallel to the principal rotation axis of the engine to opposite sides of the engine
accessories.
[0013] The one or more engine accessories may be mounted vertically beneath the core casings.
[0014] The one or more engine accessories may include an engine accessory gearbox driven
by a take-off (e.g. radial drive shaft) from the core shaft. Other of the engine accessories
can include, for example, any one or more of an electrical power generator, a fuel
pump, an oil pump, a hydraulic pump, and an engine starter motor.
[0015] The engine accessory gearbox may include a train of spur gears which transfer the
drive to other engine accessories, the spur gears being arranged in a line and having
axes of rotation which extend perpendicularly to the principal rotation axis of the
engine. For example, the spur gears of the train may be arranged in a line with their
axes of rotation extending perpendicularly to the engine's principal rotation axis.
In particular, the line may extend in a direction which is substantially parallel
with the engine axis. This is in contrast with many conventional engine accessory
gearboxes, where the train of spur gears extend around a circumferential direction
of the engine. The train of spur gears may be mounted along a central spine member,
the other engine accessories projecting from opposite sides of the spine member.
[0016] The gas turbine engine may further include a circumferential row of outer guide vanes
located in the bypass duct rearwards of the propulsive fan, the outer guide vanes
extending radially outwardly from an inner ring which defines a radially inner surface
of the bypass duct. The aerodynamic cowl can then be rearwards of the inner ring and
can include two door sections located on respective and opposite sides of the engine,
each door section being pivotably openable about an upper edge thereof to enable maintenance
access to the engine core, and the aerodynamic cowl can further include a keel beam
which extends rearwardly from the inner ring at bottom dead centre thereof to provide
latching formations for latching to a lower edge of each door section when it is closed.
Conveniently, in such an arrangement, the ventilation inlets may be located on the
radially outer surface of the keel beam. The manifold may be located on the radially
inner surface of the keel beam. The keel beam, as well as providing a convenient location
for the ventilation inlets and/or manifold, allows the lower edges of the door sections
to be located away from bottom dead centre where core engine liquids are most likely
to accumulate. Instead these liquids can be guided and drained by appropriate shaping
of the inner surface of the keel beam. Also, the keel beam can provided a relative
rigid structure for sealing to the lower edges of the door sections, thereby improving
these seals and hence the fire zone boundary performance between core and bypass zones
of the engine.
[0017] The gas turbine engine may further include a power gearbox that receives an input
from the core shaft and outputs drive to the propulsive fan so as to drive the fan
at a lower rotational speed than the core shaft.
[0018] The turbine may be a first turbine, the compressor may be a first compressor, and
the core shaft may be a first core shaft. The engine core may then further include
a second turbine, a second compressor, and a second core shaft connecting the second
turbine to the second compressor. The second turbine, second compressor, and second
core shaft may be arranged to rotate at a higher rotational speed than the first core
shaft.
[0019] As noted elsewhere herein, arrangements of the present disclosure may be particularly,
although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly,
the gas turbine engine may comprise a gearbox that receives an input from the core
shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed
than the core shaft. The input to the gearbox may be directly from the core shaft,
or indirectly from the core shaft, for example via a spur shaft and/or gear. The core
shaft may rigidly connect the turbine and the compressor, such that the turbine and
compressor rotate at the same speed (with the fan rotating at a lower speed).
[0020] The gas turbine engine as described and/or claimed herein may have any suitable general
architecture. For example, the gas turbine engine may have any desired number of shafts
that connect turbines and compressors, for example one, two or three shafts. Purely
by way of example, the turbine connected to the core shaft may be a first turbine,
the compressor connected to the core shaft may be a first compressor, and the core
shaft may be a first core shaft. The engine core may further comprise a second turbine,
a second compressor, and a second core shaft connecting the second turbine to the
second compressor. The second turbine, second compressor, and second core shaft may
be arranged to rotate at a higher rotational speed than the first core shaft.
[0021] In such an arrangement, the second compressor may be positioned axially downstream
of the first compressor. The second compressor may be arranged to receive (for example
directly receive, for example via a generally annular duct) flow from the first compressor.
[0022] The gearbox may be arranged to be driven by the core shaft that is configured to
rotate (for example in use) at the lowest rotational speed (for example the first
core shaft in the example above). For example, the gearbox may be arranged to be driven
only by the core shaft that is configured to rotate (for example in use) at the lowest
rotational speed (for example only be the first core shaft, and not the second core
shaft, in the example above). Alternatively, the gearbox may be arranged to be driven
by any one or more shafts, for example the first and/or second shafts in the example
above.
[0023] In any gas turbine engine as described and/or claimed herein, a combustor may be
provided axially downstream of the fan and compressor(s). For example, the combustor
may be directly downstream of (for example at the exit of) the second compressor,
where a second compressor is provided. By way of further example, the flow at the
exit to the combustor may be provided to the inlet of the second turbine, where a
second turbine is provided. The combustor may be provided upstream of the turbine(s).
[0024] The or each compressor (for example the first compressor and second compressor as
described above) may comprise any number of stages, for example multiple stages. Each
stage may comprise a row of rotor blades and a row of stator vanes, which may be variable
stator vanes (in that their angle of incidence may be variable). The row of rotor
blades and the row of stator vanes may be axially offset from each other.
[0025] The or each turbine (for example the first turbine and second turbine as described
above) may comprise any number of stages, for example multiple stages. Each stage
may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades
and the row of stator vanes may be axially offset from each other.
[0026] Each fan blade may be defined as having a radial span extending from a root (or hub)
at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span
position. The ratio of the radius of the fan blade at the hub to the radius of the
fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38
0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The
ratio of the radius of the fan blade at the hub to the radius of the fan blade at
the tip may be in an inclusive range bounded by any two of the values in the previous
sentence (i.e. the values may form upper or lower bounds). These ratios may commonly
be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the
tip may both be measured at the leading edge (or axially forwardmost) part of the
blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan
blade, i.e. the portion radially outside any platform.
[0027] The radius of the fan may be measured between the engine centreline and the tip of
a fan blade at its leading edge. The fan diameter (which may simply be twice the radius
of the fan) may be greater than (or on the order of) any of: 250 cm (around 100 inches),
260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115
inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around
130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm
(around 145 inches), 380 (around 150 inches) cm or 390 cm (around 155 inches). The
fan diameter may be in an inclusive range bounded by any two of the values in the
previous sentence (i.e. the values may form upper or lower bounds).
[0028] The rotational speed of the fan may vary in use. Generally, the rotational speed
is lower for fans with a higher diameter. Purely by way of non-limitative example,
the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for
example less than 2300 rpm. Purely by way of further non-limitative example, the rotational
speed of the fan at cruise conditions for an engine having a fan diameter in the range
of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be in the range of from
1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example
in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative
example, the rotational speed of the fan at cruise conditions for an engine having
a fan diameter in the range of from 320 cm to 380 cm may be in the range of from 1200
rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example
in the range of from 1400 rpm to 1600 rpm.
[0029] In use of the gas turbine engine, the fan (with associated fan blades) rotates about
a rotational axis. This rotation results in the tip of the fan blade moving with a
velocity U
tip.
[0030] The work done by the fan blades 13 on the flow results in an enthalpy rise dH of
the flow. A fan tip loading may be defined as dH/U
tip2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across
the fan and U
tip is the (translational) velocity of the fan tip, for example at the leading edge of
the tip (which may be defined as fan tip radius at leading edge multiplied by angular
speed). The fan tip loading at cruise conditions may be greater than (or on the order
of) any of: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all
units in this paragraph being Jkg
-1K
-1/(ms
-1)
2). The fan tip loading may be in an inclusive range bounded by any two of the values
in the previous sentence (i.e. the values may form upper or lower bounds).
[0031] Gas turbine engines in accordance with the present disclosure may have any desired
bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate
of the flow through the bypass duct to the mass flow rate of the flow through the
core at cruise conditions. In some arrangements the bypass ratio may be greater than
(or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5,
14, 14.5, 15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive range
bounded by any two of the values in the previous sentence (i.e. the values may form
upper or lower bounds). The bypass duct may be substantially annular. The bypass duct
may be radially outside the core engine. The radially outer surface of the bypass
duct may be defined by a nacelle and/or a fan case.
[0032] The overall pressure ratio of a gas turbine engine as described and/or claimed herein
may be defined as the ratio of the stagnation pressure upstream of the fan to the
stagnation pressure at the exit of the highest pressure compressor (before entry into
the combustor). By way of non-limitative example, the overall pressure ratio of a
gas turbine engine as described and/or claimed herein at cruise may be greater than
(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The
overall pressure ratio may be in an inclusive range bounded by any two of the values
in the previous sentence (i.e. the values may form upper or lower bounds).
[0033] Specific thrust of an engine may be defined as the net thrust of the engine divided
by the total mass flow through the engine. At cruise conditions, the specific thrust
of an engine described and/or claimed herein may be less than (or on the order of)
any of the following: 110 Nkg
-1s, 105 Nkg
-1s, 100 Nkg
-1s, 95 Nkg
-1s, 90 Nkg
-1s, 85 Nkg
-1s or 80 Nkg
-1s. The specific thrust may be in an inclusive range bounded by any two of the values
in the previous sentence (i.e. the values may form upper or lower bounds). Such engines
may be particularly efficient in comparison with conventional gas turbine engines.
[0034] A gas turbine engine as described and/or claimed herein may have any desired maximum
thrust. Purely by way of non-limitative example, a gas turbine as described and/or
claimed herein may be capable of producing a maximum thrust of at least (or on the
order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN,
400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded
by any two of the values in the previous sentence (i.e. the values may form upper
or lower bounds). The thrust referred to above may be the maximum net thrust at standard
atmospheric conditions at sea level plus 15 deg C (ambient pressure 101.3kPa, temperature
30 deg C), with the engine static.
[0035] In use, the temperature of the flow at the entry to the high pressure turbine may
be particularly high. This temperature, which may be referred to as TET, may be measured
at the exit to the combustor, for example immediately upstream of the first turbine
vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may
be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K,
1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two
of the values in the previous sentence (i.e. the values may form upper or lower bounds).
The maximum TET in use of the engine may be, for example, at least (or on the order
of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum
TET may be in an inclusive range bounded by any two of the values in the previous
sentence (i.e. the values may form upper or lower bounds). The maximum TET may occur,
for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.
[0036] A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein
may be manufactured from any suitable material or combination of materials. For example
at least a part of the fan blade and/or aerofoil may be manufactured at least in part
from a composite, for example a metal matrix composite and/or an organic matrix composite,
such as carbon fibre. By way of further example at least a part of the fan blade and/or
aerofoil may be manufactured at least in part from a metal, such as a titanium based
metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel
based material. The fan blade may comprise at least two regions manufactured using
different materials. For example, the fan blade may have a protective leading edge,
which may be manufactured using a material that is better able to resist impact (for
example from birds, ice or other material) than the rest of the blade. Such a leading
edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus,
purely by way of example, the fan blade may have a carbon-fibre or aluminium based
body (such as an aluminium lithium alloy) with a titanium leading edge.
[0037] A fan as described and/or claimed herein may comprise a central portion, from which
the fan blades may extend, for example in a radial direction. The fan blades may be
attached to the central portion in any desired manner. For example, each fan blade
may comprise a fixture which may engage a corresponding slot in the hub (or disc).
Purely by way of example, such a fixture may be in the form of a dovetail that may
slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan
blade to the hub/disc. By way of further example, the fan blades maybe formed integrally
with a central portion. Such an arrangement may be referred to as a blisk or a bling.
Any suitable method may be used to manufacture such a blisk or bling. For example,
at least a part of the fan blades may be machined from a block and/or at least part
of the fan blades may be attached to the hub/disc by welding, such as linear friction
welding.
[0038] The gas turbine engines described and/or claimed herein may or may not be provided
with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit
area of the bypass duct to be varied in use. The general principles of the present
disclosure may apply to engines with or without a VAN.
[0039] The fan of a gas turbine as described and/or claimed herein may have any desired
number of fan blades, for example 16, 18, 20, or 22 fan blades.
[0040] As used herein, cruise conditions may mean cruise conditions of an aircraft to which
the gas turbine engine is attached. Such cruise conditions may be conventionally defined
as the conditions at mid-cruise, for example the conditions experienced by the aircraft
and/or engine at the midpoint (in terms of time and/or distance) between top of climb
and start of decent.
[0041] Purely by way of example, the forward speed at the cruise condition may be any point
in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to
0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81,
for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of
from 0.8 to 0.85. Any single speed within these ranges may be the cruise condition.
For some aircraft, the cruise conditions may be outside these ranges, for example
below Mach 0.7 or above Mach 0.9.
[0042] Purely by way of example, the cruise conditions may correspond to standard atmospheric
conditions at an altitude that is in the range of from 10000m to 15000m, for example
in the range of from 10000m to 12000m, for example in the range of from 10400m to
11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example
in the range of from 10600m to 11400m, for example in the range of from 10700m (around
35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example
in the range of from 10900m to 11100m, for example on the order of 11000m. The cruise
conditions may correspond to standard atmospheric conditions at any given altitude
in these ranges.
[0043] Purely by way of example, the cruise conditions may correspond to: a forward Mach
number of 0.8; a pressure of 23000 Pa; and a temperature of -55 deg C.
[0044] As used anywhere herein, "cruise" or "cruise conditions" may mean the aerodynamic
design point. Such an aerodynamic design point (or ADP) may correspond to the conditions
(comprising, for example, one or more of the Mach Number, environmental conditions
and thrust requirement) for which the fan is designed to operate. This may mean, for
example, the conditions at which the fan (or gas turbine engine) is designed to have
optimum efficiency.
[0045] In use, a gas turbine engine described and/or claimed herein may operate at the cruise
conditions defined elsewhere herein. Such cruise conditions may be determined by the
cruise conditions (for example the mid-cruise conditions) of an aircraft to which
at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide
propulsive thrust.
[0046] The skilled person will appreciate that except where mutually exclusive, a feature
or parameter described in relation to any one of the above aspects may be applied
to any other aspect. Furthermore, except where mutually exclusive, any feature or
parameter described herein may be applied to any aspect and/or combined with any other
feature or parameter described herein.
[0047] Embodiments will now be described by way of example only, with reference to the Figures,
in which:
Figure 1 is a sectional side view of a gas turbine engine;
Figure 2 is a close up sectional side view of an upstream portion of a gas turbine
engine;
Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine;
Figure 4 shows schematically a perspective view from the rear of an engine without
its nacelle and without most of its inner cowl;
Figure 5 shows schematically at left a transverse cross-section through the engine
of Figure 4 when door sections of its inner cowl are closed, and at right a transverse
cross-section through the engine when these door sections are swung open;
Figure 6 shows a side perspective view of front end portions of a keel beam and lower
splitter of the engine of Figure 4;
Figure 7 shows a top perspective view of the front end portions of the keel beam and
lower splitter of the engine of Figure 4;
Figure 8 shows a cooling flow on a schematic side view of the engine without its nacelle
and without most of the inner cowl; and
Figure 9 shows the cooling flow on a perspective view looking down on to the keel
beam of the engine.
[0048] Figure 1 illustrates a gas turbine engine 10 having a principal rotational axis 9.
The engine is mounted to an airframe, e.g. under a wing, by a mounting pylon 46. The
engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows:
a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core
11 that receives the core airflow A. The engine core 11 comprises, in axial flow series,
a low pressure compressor 14, a high-pressure compressor 15, combustion equipment
16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle
20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22
and a bypass exhaust nozzle 18. The pylon 46 forms a bifurcation in the bypass duct
where it traverses the duct to join to the engine core 11. The bypass airflow B flows
through the bypass duct 22, where it is straightened by a row of outer guide vanes
40 before exiting the bypass exhaust nozzle 18. Rearward of the outer guide vanes
40, the engine core 10 is surrounded by an inner cowl 41 which provides an aerodynamic
fairing defining an inner surface of the bypass duct 22. The fan 23 is attached to
and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox
30.
[0049] In use, the core airflow A is accelerated and compressed by the low pressure compressor
14 and directed into the high pressure compressor 15 where further compression takes
place. The compressed air exhausted from the high pressure compressor 15 is directed
into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted.
The resultant hot combustion products then expand through, and thereby drive, the
high pressure and low pressure turbines 17, 19 before being exhausted through the
nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the
high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally
provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction
gearbox.
[0050] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure
2. The low pressure turbine 19 (see Figure 1) drives the shaft 26, which is coupled
to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly
of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that
are coupled together by a planet carrier 34. The planet carrier 34 constrains the
planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling
each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled
via linkages 36 to the fan 23 in order to drive its rotation about the engine axis
9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus
or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure
24.
[0051] Note that the terms "low pressure turbine" and "low pressure compressor" as used
herein may be taken to mean the lowest pressure turbine stages and lowest pressure
compressor stages (i.e. not including the fan 23) respectively and/or the turbine
and compressor stages that are connected together by the interconnecting shaft 26
with the lowest rotational speed in the engine (i.e. not including the gearbox output
shaft that drives the fan 23). In some literature, the "low pressure turbine" and
"low pressure compressor" referred to herein may alternatively be known as the "intermediate
pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature
is used, the fan 23 may be referred to as a first, or lowest pressure, compression
stage.
[0052] The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3.
Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their
periphery to intermesh with the other gears. However, for clarity only exemplary portions
of the teeth are illustrated in Figure 3. There are four planet gears 32 illustrated,
although it will be apparent to the skilled reader that more or fewer planet gears
32 may be provided within the scope of the claimed invention. Practical applications
of a planetary epicyclic gearbox 30 generally comprise at least three planet gears
32.
[0053] The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the
planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages
36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox
30 may be used. By way of further example, the epicyclic gearbox 30 may be a star
arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus)
gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring
gear 38. By way of further alternative example, the gearbox 30 may be a differential
gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.
[0054] It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of
example only, and various alternatives are within the scope of the present disclosure.
Purely by way of example, any suitable arrangement may be used for locating the gearbox
30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way
of further example, the connections (such as the linkages 36, 40 in the Figure 2 example)
between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26,
the output shaft and the fixed structure 24) may have any desired degree of stiffness
or flexibility. By way of further example, any suitable arrangement of the bearings
between rotating and stationary parts of the engine (for example between the input
and output shafts from the gearbox and the fixed structures, such as the gearbox casing)
may be used, and the disclosure is not limited to the exemplary arrangement of Figure
2. For example, where the gearbox 30 has a star arrangement (described above), the
skilled person would readily understand that the arrangement of output and support
linkages and bearing locations would typically be different to that shown by way of
example in Figure 2.
[0055] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement
of gearbox styles (for example star or planetary), support structures, input and output
shaft arrangement, and bearing locations.
[0056] Optionally, the gearbox may drive additional and/or alternative components (e.g.
the intermediate pressure compressor and/or a booster compressor).
[0057] Other gas turbine engines to which the present disclosure may be applied may have
alternative configurations. For example, such engines may have an alternative number
of compressors and/or turbines and/or an alternative number of interconnecting shafts.
By way of further example, the gas turbine engine shown in Figure 1 has a split flow
nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle
that is separate to and radially outside the core engine nozzle 20. However, this
is not limiting, and any aspect of the present disclosure may also apply to engines
in which the flow through the bypass duct 22 and the flow through the core 11 are
mixed, or combined, before (or upstream of) a single nozzle, which may be referred
to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have
a fixed or variable area. In some arrangements, the gas turbine engine 10 may not
comprise a gearbox 30.
[0058] The geometry of the gas turbine engine 10, and components thereof, is defined by
a conventional axis system, comprising an axial direction (which is aligned with the
rotational axis 9), a radial direction (in the bottom-to-top direction in Figure 1),
and a circumferential direction (perpendicular to the page in the Figure 1 view).
The axial, radial and circumferential directions are mutually perpendicular.
[0059] Figure 4 shows schematically a perspective view from the rear of the engine 10 with
the nacelle 21, most of the inner cowl 41 and the pylon 46 removed. A fan case 42
defines an outer surface of the bypass duct 22 and towards the rear of the fan case
an inner ring 47 defines an inner surface of the bypass duct 22. The outer guide vanes
40 extend radially from the inner ring to the fan case, and the engine core 11 projects
rearwardly from the plane of the outer guide vanes. A fixture 45 located at top dead
centre behind the inner ring provides a connection point for the mounting pylon 46
which mounts the engine to the airframe.
[0060] The inner cowl 41 includes two door sections, one on either side of the engine 10,
with each door section being pivotably openable about a respective pivot line which
extends from front to back along that door section's side of the mounting pylon 46.
This allows the door sections to be swung upwards and away from the engine core 11
for maintenance access. Conveniently, the top edges of the door sections can form
the pivot lines. Figure 5 shows schematically at left a transverse cross-section through
the engine when the door sections of the inner cowl 41 are closed (and also door sections
of the nacelle 21 surrounding the fan case 42 are closed), and at right a transverse
cross-section through the engine when these door sections are swung open.
[0061] Returning to Figure 4, a two-part inner barrel 43 attached to the inner ring 47 bridges
the space between the inner ring and the inner cowl, and provides circumferentially
extending V-grooves at its rear edge for engaging the front edges of the door sections
when they are closed. In addition, a keel beam 48 extends rearwardly from bottom dead
centre of the inner ring 47 to provide latching formations for latching to lower edges
of the door sections when they are closed. The keel beam 48 can have one or more cut-out
drain holes (not shown) to provide core zone liquid drainage, and a profiled drainage
guidance surface to provide a positive drainage force via gravity to the drain holes.
[0062] A lower splitter 49 may also traverse the bypass duct, and conveniently, the keel
beam can then form an interface for mounting a radially inner end of the lower splitter.
This mounting arrangement is advantageous when the engine has a large volume of core-mounted
accessories in the space between the engine core and the keel beam such that a direct
mount of the lower splitter to a casing of the engine core is impractical. The lower
splitter can be fastened to the keel beam by a straightforward bolting arrangement
positioned where the components are in contact.
[0063] Figures 6 and 7 show side and top perspective views of front end portions of the
keel beam 48 and lower splitter 49. They also show engine accessories 40 (drawn semi-transparently
in Figures 6 and 7) which are located above the keel beam within an engine zone bounded
on a radially outer side by the inner cowl 41, and on a radially inner side by core
casings 44 of the engine core 11. The engine accessories 40 include an accessory gearbox
driven by a take-off (such as a radial drive shaft - not shown) from the core shaft
26. Other accessories may include any one or more of a power generator, a fuel pump,
an oil pump, a hydraulic pump, and an engine starter motor. The engine accessory gearbox
has a front end that receives the drive from the drive shaft 26 and has a train of
spur gears which transfer the drive to the other accessories. These spur gears are
arranged in a line along a central spine member 52, with the other accessories projecting
from opposite sides of the spine member. The central spine member extends substantially
parallel to the engine axis 9 with the rotation axes of the spur gears perpendicular
to the engine axis.
[0064] The engine zone in which the engine accessories 40 are located can reach air temperatures
above 300°C. Thus a ventilation system is provided to maintain engine accessories
within acceptable thermal limits. This system is located on the keel beam 48 and comprises
a ventilation inlet 60 which faces forwards at engine bottom dead centre in the bypass
duct 22 to receive a portion of the relatively low temperature (<100°C) bypass airflow
B as a cooling flow for the accessories. Kinetic energy of the received bypass flow
is converted at the inlet into an increased pressure of the cooling flow, whereby
the cooling flow has a pressure differential relative to air pressure in the engine
zone which is sufficient to keep the accessories ventilated even at low engine power
conditions. To direct the cooling flow where it is needed, the ventilation system
further comprises a manifold 61 located on the radially inner surface of the keel
beam in fluid communication with the ventilation inlet to collect the cooling flow
therefrom. The cooling flow leaves the manifold through plural exhaust holes 62 to
impinge upon and extract heat from the engine accessories, protecting the accessories
from the heat present in the engine zone. For example, the manifold and its exhaust
holes can be configured as shown so that the holes form two lines which extend substantially
parallel to the engine axis 9 to opposite sides of the accessories. In this way, the
cooling air can be exhausted at high velocities over the full length of the engine
accessories 40.
[0065] Optionally, the ventilation system also comprises a control unit 63 which is under
the control of the engine's EEC, and which regulates the amount of cooling flow required
to ventilate the engine accessories 40 at different flight conditions thereby modulating
and optimising the cooling flow at different flight conditions, to reduce its impact
on engine performance and specific fuel consumption. Conveniently, the control unit
can be located between the inlet 60 and the manifold 61 to mediate transfer of the
cooling flow therebetween. However, particularly if regulation of the cooling flow
has a little impact on engine performance, the control unit can be omitted to reduce
weight and complexity of the ventilation system.
[0066] Figure 8 shows a schematic side view of the engine with the nacelle 21, most of the
inner cowl 41 and the pylon 46 removed, and Figure 9 shows a perspective view looking
down on to the keel beam 48. The bypass airflow B received into the inlet 60, and
the cooling flow through the manifold 61 and out of the exhaust holes 62 are indicated
by arrowed lines and illustrate how the cooling flow surrounds the engine accessories
40 in a continuous film of cooled air.
[0067] The ventilation inlet 60 takes advantage of the dynamic head of the bypass airflow
B to generate a sufficient driving force for the flow to travel through the control
unit 63 and the manifold 61. For example, the ventilation inlet can be a total pressure
inlet, e.g. converting at least 50% of the kinetic energy of the received bypass airflow
into a pressure rise, and preferably substantially stagnating the received bypass
airflow.
[0068] Although shown in Figures 6 to 9 with just one ventilation inlet 60, the ventilation
system may have plural inlets. Additionally, or alternatively, the inlet or inlets
can be located elsewhere, such as on the lower splitter 49.
[0069] It will be understood that the invention is not limited to the embodiments above-described
and various modifications and improvements can be made without departing from the
concepts described herein. Except where mutually exclusive, any of the features may
be employed separately or in combination with any other features and the disclosure
extends to and includes all combinations and sub-combinations of one or more features
described herein.
1. A gas turbine engine (10) for an aircraft, the engine including:
an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft
(26) connecting the turbine to the compressor;
core casings (44) surrounding the engine core;
an aerodynamic cowl (41) which surrounds the core casings (44);
a propulsive fan (23) located upstream of the engine core, the fan generating a core
airflow which enters the core engine and a bypass airflow which enters a bypass duct
(22) surrounding the aerodynamic cowl (41); and
one or more engine accessories (40) mounted in a space between the core casings (44)
and the aerodynamic cowl (41);
wherein the gas turbine engine further includes:
one or more ventilation inlets (60) which receive a portion of the bypass airflow
as a cooling flow for the engine accessories (40), the ventilation inlets being configured
to convert at least a portion of the kinetic energy of the bypass flow received therein
into a pressure rise such that the pressure of the cooling flow is increased relative
to the pressure of the bypass airflow; and
a manifold (61) in fluid communication with the ventilation inlets (60) to collect
the cooling flow therefrom, the manifold having plural exhaust holes (62) therefrom
through which the cooling flow leaves the manifold to impinge upon and thereby cool
the engine accessories (40).
2. A gas turbine engine (10) according to claim 1, wherein the one or more ventilation
inlets (60) convert at least 50% of the kinetic energy of the portion of the bypass
airflow received therein into a pressure rise.
3. A gas turbine engine (10) according to claim 1 or 2, which further includes a control
unit (63) which is operable to regulate the amount of cooling flow collected by the
manifold from the ventilation inlets.
4. A gas turbine engine (10) according to any one of the previous claims, wherein the
exhaust holes (62) are arranged such that the cooling flow impinging upon the engine
accessories (40) forms a continuous film of cooling air around the accessories.
5. A gas turbine engine (10) according to any one of the previous claims, wherein the
exhaust holes (62) are formed into two lines which extend substantially parallel to
the principal rotation axis (9) of the engine to opposite sides of the engine accessories
(40).
6. A gas turbine engine (10) according to any one of the previous claims, wherein the
one or more engine accessories (40) are mounted vertically beneath the core casings
(44).
7. A gas turbine engine (10) according to any one of the previous claims, wherein the
one or more engine accessories (40) include an engine accessory gearbox driven by
a take-off from the core shaft.
8. A gas turbine engine (10) according to claim 7, wherein the engine accessory gearbox
includes a train of spur gears which transfer the drive to other engine accessories,
the spur gears being arranged in a line and having axes of rotation which extend perpendicularly
to the principal rotation axis of the engine.
9. A gas turbine engine (10) according to claim 8, wherein the train of spur gears is
mounted along a central spine member (52), the other engine accessories projecting
from opposite sides of the spine member.
10. A gas turbine engine (10) according to any one of the previous claims, which further
includes a circumferential row of outer guide vanes (40) located in the bypass duct
(22) rearwards of the propulsive fan (23), the outer guide vanes extending radially
outwardly from an inner ring (47) which defines a radially inner surface of the bypass
duct (22);
wherein the aerodynamic cowl is rearwards of the inner ring (47) and includes two
door sections located on respective and opposite sides of the engine, each door section
being pivotably openable about an upper edge thereof to enable maintenance access
to the engine core (11), the aerodynamic cowl further including a keel beam (48) which
extends rearwardly from the inner ring (44) at bottom dead centre thereof to provide
latching formations for latching to a lower edge of each door section when it is closed;
and
wherein the ventilation inlets (60) are located on the radially outer surface of the
keel beam (48).
11. A gas turbine engine (10) according to claim 10, wherein the manifold (61) is located
on the radially inner surface of the keel beam (48).
12. A gas turbine engine (10) according to any one of the previous claims, further including
a power gearbox (30) that receives an input from the core shaft (26) and outputs drive
to the propulsive fan (23) so as to drive the fan at a lower rotational speed than
the core shaft.
13. A gas turbine engine (10) according to any one of the previous claims, wherein the
turbine is a first turbine (19), the compressor is a first compressor (14), and the
core shaft is a first core shaft (26);
the engine core further includes a second turbine (17), a second compressor (15),
and a second core shaft (27) connecting the second turbine to the second compressor;
and
wherein the second turbine, second compressor, and second core shaft are arranged
to rotate at a higher rotational speed than the first core shaft.